Low noise compressor and turbine for geared turbofan engine

ABSTRACT

A gas turbine engine has a fan, a turbine section having a first turbine including a first turbine, a compressor, and a gear reduction positioned between the fan and the first turbine. Each of the compressor and the first turbine includes a number of blades in each of a plurality of blade rows, the number of blades rotatable at least some of the time at a rotational speed in operation, and the number of blades and the rotational speed being such that the following formula holds true for the plurality of the blade rows of the first turbine, but does not hold true for any of the blade rows of the compressor rotor: (number of blades×rotational speed)/60≧5500, and the rotational speed being an approach speed in revolutions per minute.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.15/270,027, filed Sep. 20, 2016, which is a continuation of U.S. patentapplication Ser. No. 15/014,363, filed Feb. 3, 2016, which is acontinuation of U.S. patent application Ser. No. 14/967,478, filed Dec.14, 2015, which is a continuation-in-part of U.S. patent applicationSer. No. 14/591,975, filed Jan. 8, 2015, which is a continuation-in-partof U.S. patent application Ser. No. 14/144,710, filed Dec. 31, 2013,which is a continuation of U.S. patent application Ser. No. 14/016,436,filed Sep. 3, 2013, now U.S. Pat. No. 8,714,913, issued May 6, 2014,which is a continuation of U.S. patent application Ser. No. 13/630,276,filed Sep. 28, 2012, now U.S. Pat. No. 8,632,301, issued Jan. 21, 2014.

BACKGROUND

This application relates to the design of a gas turbine engine rotorwhich can be operated to produce noise that is less sensitive to humanhearing.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor anddelivered downstream into a combustor section where it was mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving the turbine rotors to rotate.

Typically, there is a high pressure turbine rotor, and a low pressureturbine rotor. Each of the turbine rotors include a number of rows ofturbine blades which rotate with the rotor. Interspersed between therows of turbine blades are vanes.

The high pressure turbine rotor has typically driven a high pressurecompressor rotor, and the low pressure turbine rotor has typicallydriven a low pressure compressor rotor. Each of the compressor rotorsalso include a number of compressor blades which rotate with the rotors.There are also vanes interspersed between the rows of compressor blades.

The low pressure turbine or compressor can be a significant noisesource, as noise is produced by fluid dynamic interaction between theblade rows and the vane rows. These interactions produce tones at ablade passage frequency of each of the low pressure turbine rotors, thelow pressure compressor rotors, and their harmonics.

The noise can often be in a frequency range that is very sensitive tohumans. To mitigate this problem, in the past, a vane-to-blade ratio hasbeen controlled to be above a certain number. As an example, avane-to-blade ratio may be selected to be 1.5 or greater, to prevent afundamental blade passage tone from propagating to the far field. Thisis known as “cut-off.”

However, acoustically cut-off designs may come at the expense ofincreased weight and reduced aerodynamic efficiency. Stated another way,by limiting the designer to a particular vane to blade ratio, thedesigner may be restricted from selecting such a ratio based upon othercharacteristics of the intended engine.

Historically, the low pressure turbine has driven both a low pressurecompressor section and a fan section. More recently, a gear reductionhas been provided such that the fan and low pressure compressor can bedriven at distinct speeds.

SUMMARY

A gas turbine engine according to an example of the present disclosureincludes a fan section that has a fan. The fan includes at least one fanblade, and a low fan pressure ratio less than about 1.45. The low fanpressure ratio is measured across a fan blade alone. A compressor isrotationally coupled to the fan. A turbine section has a first turbineand a second turbine. A gear reduction is positioned between the firstturbine on one side and the compressor and fan on another side. The gearreduction includes an epicycle gear train that has a gear reductionratio of greater than 2.5:1, and the first turbine is positioned in adriving relationship with the gear reduction. The first turbine includesa pressure ratio greater than about 5:1. The first turbine has an inletthat has an inlet pressure, and an outlet that is prior to any exhaustnozzle and having an outlet pressure, and the pressure ratio of thefirst turbine is a ratio of the inlet pressure to the outlet pressure.Each of the compressor and the first turbine includes a number of bladesin each of a plurality of blade rows. The number of blades is rotatableat least some of the time at a rotational speed in operation, and thenumber of blades and the rotational speed is such that the followingformula holds true for more than one of the blade rows of the firstturbine, but does not hold true for any of the blade rows of thecompressor: (number of blades×rotational speed)/60≧5500, the rotationalspeed being an approach speed in revolutions per minute, taken at anapproach certification point as defined in Part 36 of the FederalAirworthiness Regulations. The gas turbine engine is rated to produce15,000 pounds of thrust or more.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number less than or equal to about 10000 for at least oneof the blade rows of the first turbine, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations.

In a further embodiment of any of the foregoing embodiments, thefollowing formula holds true for at least one of the blade rows of thefirst turbine: (number of blades×rotational speed)/60≦10000, therotational speed being a cruise speed in revolutions per minute.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number less than or equal to 7000 Hz for at least one ofthe blade rows of the first turbine, taken at an approach certificationpoint as defined in Part 36 of the Federal Airworthiness Regulations.

A further embodiment of any of the foregoing embodiments includes abypass ratio greater than ten (10).

In a further embodiment of any of the foregoing embodiments, the formularesults in a number greater than 6000 for at least one of the blade rowsof the first turbine, taken at an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations.

In a further embodiment of any of the foregoing embodiments, the fan hasa low corrected fan tip speed less than 1150 ft/second. The lowcorrected fan tip speed is an actual fan tip speed in ft/second at anambient temperature divided by [(Tambient° R)/(518.7° R)]0.5.

A further embodiment of any of the foregoing embodiments includes a coreflowpath and a mid-turbine frame having airfoils positioned in the coreflowpath. The mid-turbine frame supports at least one bearing system.

In a further embodiment of any of the foregoing embodiments, the secondturbine has two stages, and the epicyclic gear train is a planetary gearsystem.

In a further embodiment of any of the foregoing embodiments, thefollowing formula holds true for at least one of the blade rows of thefirst turbine: (number of blades×rotational speed)/60≧5500, therotational speed being a cruise speed in revolutions per minute.

A gas turbine engine according to an example of the present disclosureincludes a fan, a turbine section that has a first turbine, acompressor, and a gear reduction positioned between the fan relative toan input speed from the first turbine. Each of the compressor and thefirst turbine includes a number of blades in each of a plurality ofblade rows. The number of blades rotatable at least some of the time ata rotational speed in operation, and the number of blades and therotational speed is such that the following formula holds true for amajority of the blade rows of the first turbine, and at least one of theblade rows of the compressor: (number of blades×rotationalspeed)/60≧5500, the rotational speed being an approach speed inrevolutions per minute, taken at an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations. The engineis rated to produce 15,000 pounds of thrust or more.

In a further embodiment of any of the foregoing embodiments, the formulaholds true for less than half of the blade rows of the compressor.

In a further embodiment of any of the foregoing embodiments, the formulaholds true for all of the blade rows of the first turbine.

In a further embodiment of any of the foregoing embodiments, thefollowing formula holds true for at least one of the blade rows of thefirst turbine: (number of blades×rotational speed)/60≧5500. Therotational speed is a cruise speed in revolutions per minute.

A further embodiment of any of the foregoing embodiments includes abypass ratio greater than ten (10). The gear reduction has an epicyclegear train that has a gear reduction ratio of greater than 2.5:1, andthe first turbine is in a driving relationship with the compressor.

A further embodiment of any of the foregoing embodiments includes a lowfan pressure ratio less than 1.45, wherein the low fan pressure ratio ismeasured across a fan blade alone.

In a further embodiment of any of the foregoing embodiments, the firstturbine includes a pressure ratio greater than 5:1. The first turbineincludes an inlet having an inlet pressure, and an outlet that is priorto any exhaust nozzle that an outlet pressure, and the pressure ratio ofthe first turbine is a ratio of the inlet pressure to the outletpressure.

In a further embodiment of any of the foregoing embodiments, the fan hasa low corrected fan tip speed less than 1150 ft/second. The lowcorrected fan tip speed is an actual fan tip speed in ft/second at anambient temperature divided by [(Tambient° R)/(518.7° R)]0.5.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number greater than or equal to 5500 but less than or equalto 10000 for at least one of the blade rows of the compressor, taken atan approach certification point as defined in Part 36 of the FederalAirworthiness Regulations.

In a further embodiment of any of the foregoing embodiments, thecompressor and the fan rotate at a common speed in operation.

A further embodiment of any of the foregoing embodiments includes a coreflowpath and a mid-turbine frame that has airfoils positioned in thecore flowpath.

In a further embodiment of any of the foregoing embodiments, the formulaholds true for all of the blade rows of the first turbine, and resultsin a number greater than or equal to 5500 but less than or equal to 7000for at least one of the blade rows of the compressor, taken at anapproach certification point as defined in Part 36 of the FederalAirworthiness Regulations.

A gas turbine engine according to an example of the present disclosureincludes a fan, a turbine section that has a first turbine, acompressor, and a gear reduction positioned between the fan relative toan input speed from the first turbine. Each of the compressor and thefirst turbine includes a number of blades in each of a plurality ofblade rows. The number of blades rotatable at least some of the time ata rotational speed in operation, and the number of blades and therotational speed is such that the following formula holds true for allof the blade rows of the first turbine, and all of the blade rows of thecompressor: (number of blades×rotational speed)/60≧5500, the rotationalspeed being an approach speed in revolutions per minute, taken at anapproach certification point as defined in Part 36 of the FederalAirworthiness Regulations.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number less than or equal to about 10000 for at least oneof the blade rows of the first turbine, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations.

In a further embodiment of any of the foregoing embodiments, the firstturbine includes a pressure ratio greater than 5:1. The first turbinehas an inlet that has an inlet pressure, and an outlet that is prior toany exhaust nozzle and having an outlet pressure, and the pressure ratioof the first turbine is a ratio of the inlet pressure to the outletpressure.

A further embodiment of any of the foregoing embodiments includes a lowfan pressure ratio less than about 1.45. The low fan pressure ratio ismeasured across a fan blade alone.

In a further embodiment of any of the foregoing embodiments, the gearreduction has a gear reduction ratio of greater than about 2.5:1.

In a further embodiment of any of the foregoing embodiments, the formularesults in a number greater than or equal to 5500 but less than or equalto about 7000 for at least one of the blade rows of the compressor,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations.

In a further embodiment of any of the foregoing embodiments, the gasturbine engine is rated to produce 15,000 pounds of thrust or more. Thefan has a low corrected fan tip speed less than about 1150 ft/second.The low corrected fan tip speed is an actual fan tip speed in ft/secondat an ambient temperature divided by [(Tambient° R)/(518.7° R)]0.5.

In a further embodiment of any of the foregoing embodiments, thefollowing formula holds true for at least one of the blade rows of thefirst turbine: (number of blades×rotational speed)/60≧5500, therotational speed being a cruise speed in revolutions per minute.

In a featured embodiment, a gas turbine engine has a fan, a compressorsection having a low pressure portion and a high pressure portion, acombustor section, and a turbine having a first turbine rotor. The firstturbine rotor drives the fan. A gear reduction effects a reduction inthe speed of the fan relative to a speed of the first turbine rotor.Each of the compressor rotor and the first turbine rotor includes anumber of blades in each of a plurality of rows. The blades operate atleast some of the time at a rotational speed. The number of blades andthe rotational speed are such that the following formula holds true forat least one of the blade rows of the first turbine rotor and/or thecompressor rotor: (number of blades×rotational speed)/60≧5500, therotational speed being an approach speed in revolutions per minute.

These and other features of this application will be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

FIG. 2 shows another embodiment.

FIG. 3 shows yet another embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown), or an intermediate spool,among other systems or features. The fan section 22 drives air along abypass flowpath B in a bypass duct defined within a nacelle 15, whilethe compressor section 24 drives air along a core flowpath C forcompression and communication into the combustor section 26 thenexpansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The terms “low” and “high” as applied to speed or pressure for thespools, compressors and turbines are of course relative to each other.That is, the low speed spool operates at a lower speed than the highspeed spool, and the low pressure sections operate at lower pressurethan the high pressures sections.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. In some embodiments, the bypass ratio isless than about thirty (30), or more narrowly less than about twenty(20). In embodiments, the gear reduction ratio is less than about 5.0,or less than about 4.0. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a geared architectureengine and that the present invention is applicable to other gas turbineengines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient°R)/(518.7)° R]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second.

The use of the gear reduction between the low speed spool and the fanallows an increase of speed to the low pressure compressor. In the past,the speed of the low pressure turbine and compressor has been somewhatlimited in that the fan speed cannot be unduly large. The maximum fanspeed is at its outer tip, and in larger engines, the fan diameter ismuch larger than it may be in smaller power engines. However, the use ofthe gear reduction has freed the designer from limitation on the lowpressure turbine and compressor speeds caused by a desire to not haveunduly high fan speeds.

It has been discovered that a careful design between the number ofrotating blades, and the rotational speed of the low pressure turbinecan be selected to result in noise frequencies that are less sensitiveto human hearing. The same is true for the low pressure compressor 44.

A formula has been developed as follows:

(blade count×rotational speed)/60 sec≧5500 Hz.

That is, the number of rotating blades in any low pressure turbinestage, multiplied by the rotational speed of the low pressure turbine 46(in revolutions per minute), divided by 60 sec should be greater than orequal to about 5500 Hz. The same holds true for the low pressurecompressor stages. More narrowly, the amounts should be greater than orequal to about 6000 Hz. In embodiments, the amount is less than or equalto about 10000 Hz, or more narrowly less than or equal to about 7000 Hz.A worker of ordinary skill in the art would recognize that the 60 secfactor is to change revolutions per minute to Hertz, or revolutions perone second. For the purposes of this disclosure, the term “about” means±3% of the respective quantity unless otherwise disclosed.

The operational speed of the low pressure turbine 46 and low pressurecompressor 44 as utilized in the formula should correspond to the engineoperating conditions at each noise certification point defined in Part36 or the Federal Airworthiness Regulations. More particularly, therotational speed may be taken as an approach certification point asdefined in Part 36 of the Federal Airworthiness Regulations. Forpurposes of this application and its claims, the term “approach speed”equates to this certification point. In other embodiments, therotational speed is taken as a takeoff or cruise certification point,with the terms “takeoff speed” and “cruise speed” equating to thesecertification points. In some embodiments, the above formula results ina number that is less than or equal to about 10000 Hz at takeoff speed.In other embodiments, the above formula results in a number that is lessthan or equal to about 7000 Hz at approach speed.

It is envisioned that all of the rows in the low pressure turbine 46meet the above formula. However, this application may also extend to lowpressure turbines wherein the majority of the blade rows, or at leasthalf of the blade rows, in the low pressure turbine meet the aboveformula, but perhaps some may not. By implication at least one, or lessthan half, of the rows meet the formula. The same is true for lowpressure compressors, wherein all of the rows in the low pressurecompressor 44 would meet the above formula. However, the application mayextend to low pressure compressors wherein only the majority of theblade rows, or at least half of the blade rows, in the low pressurecompressor meet the above formula, but some perhaps may not. Of course,by implication the formula may be true for at least some of the turbinerows but no compressor rows. In some cases, only one row of the lowpressure turbine and/or low pressure compressor may meet the formula.Also, the formula may apply to at least some compressor rows, but no rowin the turbine meets the formula.

This will result in operational noise that would be less sensitive tohuman hearing.

In embodiments, it may be that the formula can result in a range ofgreater than or equal to 5500 Hz, and moving higher. Thus, by carefullydesigning the number of blades and controlling the operational speed ofthe low pressure turbine 46 (and a worker of ordinary skill in the artwould recognize how to control this speed) one can assure that the noisefrequencies produced by the low pressure turbine are of less concern tohumans.

The same holds true for designing the number of blades and controllingthe speed of the low pressure compressor 44. Again, a worker of ordinaryskill in the art would recognize how to control the speed.

In embodiments, it may be only the low pressure turbine rotor 46, or thelow pressure compressor rotor 44 which is designed to meet the meet theabove formula. On the other hand, it is also possible to ensure thatboth the low pressure turbine 46 and low pressure compressor 44 meet theabove formula.

This invention is most applicable to jet engines rated to produce 15,000pounds of thrust or more. In this thrust range, prior art jet engineshave typically had frequency ranges of about 4000 hertz. Thus, the noiseproblems as mentioned above have existed.

Lower thrust engines (<15,000 pounds) may have operated under conditionsthat sometimes passed above the 4000 Hz number, and even approached 6000Hz, however, this has not been in combination with the gearedarchitecture, nor in the higher powered engines which have the largerfans, and thus the greater limitations on low pressure turbine or lowpressure compressor speed.

FIG. 2 shows an embodiment 200, wherein there is a fan drive turbine 208driving a shaft 206 to in turn drive a fan rotor 202. A gear reduction204 may be positioned between the fan drive turbine 208 and the fanrotor 202. This gear reduction 204 may be structured and operate likethe gear reduction disclosed above. A compressor rotor 210 is driven byan intermediate pressure turbine 212, and a second stage compressorrotor 214 is driven by a turbine rotor 216. A combustion section 218 ispositioned intermediate the compressor rotor 214 and the turbine section216.

FIG. 3 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a low pressureturbine section.

The FIGS. 2 and 3 engines may be utilized with the speed and bladefeatures disclosed above.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

1. A gas turbine engine comprising: a fan section including a fan, thefan including at least one fan blade; a low fan pressure ratio less thanabout 1.45, wherein the low fan pressure ratio is measured across a fanblade alone; a compressor rotationally coupled to the fan; a turbinesection having a first turbine and a second turbine; a gear reductionpositioned between the first turbine on one side and the compressor andfan on another side, the gear reduction including an epicycle gear trainhaving a gear reduction ratio of greater than 2.5:1, and the firstturbine positioned in a driving relationship with the gear reduction;the first turbine including a pressure ratio greater than about 5:1, thefirst turbine including an inlet having an inlet pressure, and an outletthat is prior to any exhaust nozzle and having an outlet pressure, andthe pressure ratio of the first turbine being a ratio of the inletpressure to the outlet pressure; each of the compressor and the firstturbine includes a number of blades in each of a plurality of bladerows, the number of blades rotatable at least some of the time at arotational speed in operation, and the number of blades and therotational speed being such that the following formula holds true formore than one of the blade rows of the first turbine, but does not holdtrue for any of the blade rows of the compressor:(number of blades×rotational speed)/60≧5500, and the rotational speedbeing an approach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations; and wherein the gas turbine engine is rated to produce15,000 pounds of thrust or more.
 2. The gas turbine engine as set forthin claim 1, wherein the formula results in a number less than or equalto about 10000 for at least one of the blade rows of the first turbine,taken at an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations.
 3. The gas turbine engine as setforth in claim 2, wherein the following formula holds true for at leastone of the blade rows of the first turbine:(number of blades×rotational speed)/60≦10000, the rotational speed beinga cruise speed in revolutions per minute.
 4. The gas turbine engine asset forth in claim 2, wherein the formula results in a number less thanor equal to 7000 Hz for at least one of the blade rows of the firstturbine, taken at an approach certification point as defined in Part 36of the Federal Airworthiness Regulations.
 5. The gas turbine engine asset forth in claim 1, further comprising a bypass ratio greater than ten(10).
 6. The gas turbine engine as set forth in claim 5, wherein theformula results in a number greater than 6000 for at least one of theblade rows of the first turbine, taken at an approach certificationpoint as defined in Part 36 of the Federal Airworthiness Regulations. 7.The gas turbine engine as set forth in claim 6, wherein the fan has alow corrected fan tip speed less than 1150 ft/second, wherein the lowcorrected fan tip speed is an actual fan tip speed in ft/second at anambient temperature divided by [(Tambient° R)/(518.7° R)]^(0.5).
 8. Thegas turbine engine as set forth in claim 7, further comprising a coreflowpath and a mid-turbine frame having airfoils positioned in the coreflowpath, the mid-turbine frame supporting at least one bearing system.9. The gas turbine engine as set forth in claim 8, wherein the secondturbine has two stages, and the epicyclic gear train is a planetary gearsystem.
 10. The gas turbine engine as set forth in claim 9, wherein thefollowing formula holds true for at least one of the blade rows of thefirst turbine:(number of blades×rotational speed)/60≧5500, the rotational speed beinga cruise speed in revolutions per minute.
 11. A gas turbine enginecomprising: a fan; a turbine section including a first turbine; acompressor; a gear reduction positioned between the fan relative to aninput speed from the first turbine; and wherein each of the compressorand the first turbine includes a number of blades in each of a pluralityof blade rows, the number of blades rotatable at least some of the timeat a rotational speed in operation, and the number of blades and therotational speed being such that the following formula holds true for amajority of the blade rows of the first turbine, and at least one of theblade rows of the compressor:(number of blades×rotational speed)/60≧5500, and the rotational speedbeing an approach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations, and the engine is rated to produce 15,000 pounds of thrustor more.
 12. The gas turbine engine as set forth in claim 11, whereinthe formula holds true for less than half of the blade rows of thecompressor.
 13. The gas turbine engine as set forth in claim 12, whereinthe formula holds true for all of the blade rows of the first turbine.14. The gas turbine engine as set forth in claim 11, wherein thefollowing formula holds true for at least one of the blade rows of thefirst turbine:(number of blades×rotational speed)/60≧5500, the rotational speed beinga cruise speed in revolutions per minute.
 15. The gas turbine engine asset forth in claim 11, further comprising a bypass ratio greater thanten (10), the gear reduction includes an epicycle gear train having agear reduction ratio of greater than 2.5:1, and the first turbine is ina driving relationship with the compressor.
 16. The gas turbine engineas set forth in claim 15, further comprising a low fan pressure ratioless than 1.45, wherein the low fan pressure ratio is measured across afan blade alone.
 17. The gas turbine engine as set forth in claim 16,wherein the first turbine includes a pressure ratio greater than 5:1,the first turbine including an inlet having an inlet pressure, and anoutlet that is prior to any exhaust nozzle and having an outletpressure, and the pressure ratio of the first turbine being a ratio ofthe inlet pressure to the outlet pressure.
 18. The gas turbine engine asset forth in claim 17, wherein the fan has a low corrected fan tip speedless than 1150 ft/second, wherein the low corrected fan tip speed is anactual fan tip speed in ft/second at an ambient temperature divided by[(Tambient° R)/(518.7° R)]^(0.5).
 19. The gas turbine engine as setforth in claim 17, wherein the formula results in a number greater thanor equal to 5500 but less than or equal to 10000 for at least one of theblade rows of the compressor, taken at an approach certification pointas defined in Part 36 of the Federal Airworthiness Regulations.
 20. Thegas turbine engine as set forth in claim 19, wherein the compressor andthe fan rotate at a common speed in operation.
 21. The gas turbineengine as set forth in claim 20, further comprising a core flowpath anda mid-turbine frame having airfoils positioned in the core flowpath. 22.The gas turbine engine as set forth in claim 11, wherein the formulaholds true for all of the blade rows of the first turbine, and resultsin a number greater than or equal to 5500 but less than or equal to 7000for at least one of the blade rows of the compressor, taken at anapproach certification point as defined in Part 36 of the FederalAirworthiness Regulations.
 23. A gas turbine engine comprising: a fan; aturbine section including a first turbine; a compressor; a gearreduction positioned between the fan relative to an input speed from thefirst turbine; and wherein each of the compressor and the first turbineincludes a number of blades in each of a plurality of blade rows, thenumber of blades rotatable at least some of the time at a rotationalspeed in operation, and the number of blades and the rotational speedbeing such that the following formula holds true for all of the bladerows of the first turbine, and all of the blade rows of the compressor:(number of blades×rotational speed)/60≧5500, and the rotational speedbeing an approach speed in revolutions per minute, taken at an approachcertification point as defined in Part 36 of the Federal AirworthinessRegulations.
 24. The gas turbine engine as set forth in claim 23,wherein the formula results in a number less than or equal to about10000 for at least one of the blade rows of the first turbine, taken atan approach certification point as defined in Part 36 of the FederalAirworthiness Regulations.
 25. The gas turbine engine as set forth inclaim 23, wherein the first turbine includes a pressure ratio greaterthan 5:1, the first turbine including an inlet having an inlet pressure,and an outlet that is prior to any exhaust nozzle and having an outletpressure, and the pressure ratio of the first turbine being a ratio ofthe inlet pressure to the outlet pressure.
 26. The gas turbine engine asset forth in claim 25, further comprising a low fan pressure ratio lessthan about 1.45, wherein the low fan pressure ratio is measured across afan blade alone.
 27. The gas turbine engine as set forth in claim 26,wherein the gear reduction has a gear reduction ratio of greater thanabout 2.5:1.
 28. The gas turbine engine as set forth in claim 27,wherein the formula results in a number greater than or equal to 5500but less than or equal to about 7000 for at least one of the blade rowsof the compressor, taken at an approach certification point as definedin Part 36 of the Federal Airworthiness Regulations.
 29. The gas turbineengine as set forth in claim 23, wherein the gas turbine engine is ratedto produce 15,000 pounds of thrust or more, and wherein the fan has alow corrected fan tip speed less than about 1150 ft/second, wherein thelow corrected fan tip speed is an actual fan tip speed in ft/second atan ambient temperature divided by [(Tambient° R)/(518.7° R)]^(0.5). 30.The gas turbine engine as set forth in claim 23, wherein the followingformula holds true for at least one of the blade rows of the firstturbine:(number of blades×rotational speed)/60≧5500, the rotational speed beinga cruise speed in revolutions per minute.